This invention relates generally to turbomachinery compressors and more particularly relates to variable stator elements of such compressors.
A gas turbine engine includes, in serial flow communication, a compressor, a combustor, and turbine collectively constituting a turbomachinery core. The core is operable in a known manner, in conjunction with other engine components, to perform useful work such as providing propulsive thrust or mechanical work. One known configuration of the compressor is the “axi-centrifugal” compressor which includes an upstream axial-flow portion that discharges through a duct to a downstream centrifugal-flow portion.
Because of cycle demands for high pressure ratio and high efficiency, the diffuser throat flow area at the downstream end of the axi-centrifugal compressor is often relatively small. The result is that at part-speed conditions the maximum flow of the centrifugal portion of the compressor is less than the minimum stable flow of the axial compressor (i.e. the axial is pushed beyond its stall line). In prior art designs, compressor stability is achieved by extracting or “bleeding” excess flow from a location between the axial and the centrifugal compressor portions, or by using inlet guide vanes with a variable effective angle in the axial portion in order to throttle mass flow. Both of these methods can have negative impacts on performance and have limits to their effectiveness.
Accordingly, there remains a need for an axi-centrifugal compressor that is operable with good efficiency at varying flow rates.